Missile guidance and control has received considerable attention in the last 50 years. Proportional Navigation (PN) and its multiple variants has been the preferred guidance technique. It is unfortunately based on the following assumptions:                (i) The target has zero acceleration;        (ii) The interceptor has a perfect dynamic response and perfect control of its acceleration;        (iii) The interceptor is launched on a near collision condition;        (iv) The interceptor has no longitudinal acceleration;One or more of said assumptions is not applicable in challenging interception situations.        
U.S. Pat. No. 6,244,536 B1 uses a modified formulation of PN called PRONAV and Riccati Equation. While the use of Riccati equation reduces the divert effort, it is predicated on the availability of good state models. U.S. Pat. No. 7,185,844 B2 and also Ref 1 introduces PN with a so called “parallel” navigation additional term proportional to the cube of the line-of-sight rate and additional term proportional to relative longitudinal acceleration, not easy to estimate very accurately and account for interceptor acceleration. The compensation for such effects is beneficial. Its explicit compensation as proposed in U.S. Pat. No. 7,185,844 B2 is only as good as the estimation/measurement of corresponding terms. In the present invention compensates for such effects without complex modeling of the effects and without being subjected to the effects of possible estimation errors. A number of works have addressed the problem of highly maneuvering targets and proposed solutions that require additional information. The so-called optimal law of reference 2 is applicable to first order interceptor response, requires good estimation of target maneuver and of tgo (time to go). Neoclassical Guidance in reference 3 does not require estimating target acceleration and tgo, but since it is based on adjoint techniques, it requires a good dynamical model of the interceptor flight control system and does not apply to the case of dual concurrent lateral controls. The claim that Neoclassical Guidance can force Zero Miss distance (ZMD) against any bounded target maneuver is questioned in Ref. 4. Reference 5 includes in the observer a set of maneuver patterns that enhance the prediction of target kinematics and a method based on differential game theory to adjust the gains of the guidance. The degree to which actual target maneuvers must “match” the models makes this approach not very realistic. The quest for better prediction of target maneuvers in Ref. 6 leads to use banks of filters with typical maneuvers and maneuver detectors. Here the problem is that the detection of a maneuver of change thereof takes some time if one wants to have enough confidence in the decision made and also the “mathematical” separation of the maneuvers may not be evident. The use of a Kalman Filter transition matrix with a ZEM based guidance in Ref. 7 helps in reducing the effects of delays but is only applicable for longer range exo-atmospheric guidance.
Kalman Filtering has been a technique of choice for estimating target motion. While it produces good estimations, its relatively slow convergence may cause it to be ineffective at the end-game when rapid target maneuvers are encountered. Reference 8 showed that Higher Order Sliding Mode Observers can provide faster and yet more accurate estimations than traditional Kalman Filter.
To be effective, a guidance law must be supported by a good autopilot. Most autopilot design are based on classical control techniques or on state feedback techniques that rely on internal mathematical models are only as good as the internal models. Further, the internal models are more and more expensive and difficult to develop as the domain of utilization of the missiles becomes larger and larger and as the accuracy is more than often questionable, which then degrades the accuracy of the guidance and control (G&C). Their validation using wind tunnels is also becoming increasingly difficult due to the cost of the energy required for simulating missile flights in the atmosphere at several thousand of meters per second. Most of the numerical codes calculating the flow around the missile and most of the wind tunnel experiments simulate steady state conditions that is, their governing equations do not include or model partial derivatives with respect to time. This assumption was reasonable up to now but is becoming increasingly questionable as missile with greater and greater agility and shorter time constants are designed. Working with non steady state computational fluid codes of designing non-steady state wind tunnel experiments increases dramatically the difficulty in generating realistic missile models. Thus, there is a need for more robust control systems and methods that are tolerant of complex and unpredictable interactions and adapt to changing dynamics resulting from effects such as hypersonic aerodynamics and interactions. Such hypersonic effects are described in Ref. 9. Such robust control must be able to work with nonlinear thrusters, to be insensitive to model variations caused by variations of altitude, mass, center of gravity, and other effects difficult to model or predict and that by not being dependent of missile specific mathematical model to allow a greater reuse of previous designs on new missiles.
U.S. Pat. No. 6,532,454, B1 uses adaptive techniques to estimate an interceptor model. While this approach works well when given enough time for the estimation process to settle, it cannot handle the case where rapid disturbances are created by the interaction between the firing of thrusters and the aerodynamic flow around the missile as described in Ref. 9. Another important issue is to achieve interceptor maneuverability as large as possible. This calls for operations in the endo-atmospheric domain for combined operation of several divert mechanisms, possibly several control surfaces. U.S. Pat. RE37,331 E where a forward placed thruster and tail (fin control) are used jointly to initiate a lateral maneuver faster and without non-minimum phase effects. One of the potential problems associated with this approach is that it requires accurate estimation of lateral divert and angular motion effects of each control that are by definition not measurable separately. A similar approach is described in Reference 10 where a dual control missile is fitted with the canard, primarily used to achieve the miss-distance and the tail for “trimming” the pitch attitude; the two surfaces are controlled with first order sliding mode controls approximated by high gain saturation which unfortunately cause the controllers to lose some of their robustness. U.S. Pat. No. 6,341,249 control satellite attitude using a Lyapunov Controller or a first order sliding mode controller. A short fall associated with the use of first order sliding mode control is that it achieves its robustness by high frequency (infinite switching rate) or eventually can be approximated by a sub-performing high gain approximation. Reference 11 proposes a Multiple Input Multiple Output solution to steer the angle of attack and pitch rate to satisfy a Lateral acceleration condition. It uses a Variable Structure Solution; unfortunately, it uses Euler Angles and its generalization to a Quaternion solution would be very difficult. U.S. Pat. Nos. 6,611,823 B1 and 7,080,055 B2 both address the problems associated with actuator nonlinearities as encountered with on-off thrusters and come up with complex neural network solutions requiring some “learning” hardly an option with missile systems.
Thus, there is a need for more robust control systems and methods that are tolerant of complex and unpredictable interactions and adapt to changing dynamics resulting from effects such as hypersonic aerodynamics and interactions